Gas turbine engine blades typically have dovetails or roots carried by a slot in a metal rotor disk or drum rotor. A typical blade 1 is shown in FIG. 1 with an airfoil section 2 and a root section 3. The root section 3 provides the means by which the blade is attached and secured to the rotor disk or other similar component of a gas turbine engine or compressor of a turbomachine. The blade 1 may also include an interface 4 between the airfoil 2 and the root 3 to conform to the rotor disk or other attachment mechanism.
Composite laminated blades have many advantages over blades made with other materials, such as current metal alloys. They have a high strength to weight ratio that allows for the design of low weight parts that can withstand the extreme temperatures and loading of turbomachinery. They can also be designed with parts with design features not possible with other materials (such as extreme forward sweep of compressor blading). A major drawback of composite blades is their strength is essentially unidirectional. Despite having a relatively high uniaxial tensile strength, the composite materials are fragile and weak under compression or shear. However, in gas turbines, the blades are usually under extremely high tensile loads due to high rotational speeds of the rotor disk and blades. Problems usually arise with regard to the transfer of such loads into the disk. Since the blades are often made of a metal, the transfer of loads between the two can lead to damage of the fibers, or even worse, delamination of the blade material.
FIGS. 2a-2c show the problem discussed above, where there are shown three separate views of an example of a composite laminated blade root. FIG. 2a shows an unloaded blade 10a. FIG. 2b shows a blade having a tensile load T applied thereto, where the shear stress has caused a failure in the root section of the blade. FIG. 2c shows a loaded blade where the resulting stress from the tensile load T as applied to the blade from the surrounding disk cavity (not shown) has caused a delamination of the blade. The challenge therefore is to provide an optimum load path between the laminated blade and the surrounding disk.
Previously, one of the technology bathers for high performance composite laminated blades has been to provide an attachment scheme that would utilize the strength of composite materials to prevent the failure illustrated in FIGS. 2b-2c. As demonstrated in FIGS. 2a-2b, a critical important area is the blade attachment region or “neck” portion 11 of the blade, where the thicker root transitions out of the relatively thin airfoil section above the neck and root portions. This critical area is where the laminates of the airfoil portion of the blade that make up the pressure side and the suction side will diverge from each other and wrap around or encircle an insert to form the root portion of the blade. It is this portion which tends to delaminate or otherwise fail when the blade is loaded and the resulting stresses are applied to the root and interface between the root and disk. One reason for such failure is that the disk lugs tend to separate due to both the centrifugal force acting on the disk and blade due to high rotational speeds. FIG. 2d shows a blade 15 inserted into a disk 16 and under no loading from rotation. The disk lugs 17 around the neck 18 of the blade 15 define a gap G0 that conforms to the shape of the blade 15. In FIG. 2e, the blade of FIG. 2d is shown under centrifugal loading, where the gap has increased in size to GL. Although this geometrical change in the disk geometry is slight (the dimensions portrayed in FIGS. 2d-2e is exaggerated for effect), it no longer conforms to the shape of the blade. The effect of this slight increase in gap induces transverse tension and/or shear stresses in the blade as a result of the laminate in the blade conforming the a new shape of the slot formed in the disk due to the lugs 17 bending outward and increasing the gap.
Since composite laminated materials have little ability to handle transverse tension or shear loading, this will result in failure of the composite blade as in blade 10c once the intralaminar tension or shear stresses exceed the ultimate intralaminar stress capabilities of the composite material. An example would be unidirectional Kevlar composite having an ultimate intraliminar stress capability of about 6 ksi.
Also, since composite blades are very useful in a gas turbine engine, it is desirable to provide a tailored attachment mechanism of composite airfoils that both take advantage of the relatively high tensile strength of composite materials and minimizes the disadvantage of relatively low shear and transverse tension of the composite material.
U.S. Pat. No. 5,292,231 issued to Lauzeille shows a turbomachine blade made of composite laminated material, and includes a jacket wrapped around a teardrop shaped root portion. However, the jacket does not extend far along the airfoil portion of the blade to provide a compressive force against the laminates at the critical point (the point shown in FIG. 1 where the laminates digress to pass around the insert member 11). Further, the jacket does not include a thicker portion adjacent to the critical point to produce a compressive force against the laminates due to high centrifugal force acting on the blade.